A typical gas turbine engine has a flow path extending about a longitudinal axis and includes a compressor, combustor and turbine spaced sequentially along the flow path. Both the compressor and turbine include adjacent arrays of airfoils that engage fluid flowing through the flow path. The arrays are made up of rotating blades and stationary vanes. The rotating blades either transfer energy to the fluid, as in the compressor, or remove energy from the fluid, as in the turbine. Each array of vanes is located upstream of an array of blades and is configured to orient the flow of fluid for optimal engagement with the downstream blade.
In addition to the vanes, inner and outer surfaces are used to confine the flow of fluid within the annular flow path through the gas turbine engine. For the vanes, the flow surfaces are provided by platforms that are integral to the inner and outer ends of the vane. For the blades, the inner surface is provided by a platform that is integral to the blade and the outer surface is provided by a shroud having a circumferential flow surface radially outward of the tips of the blades.
The blade arrays and vane arrays are axially spaced a finite distance as a result of having adjacent rotating blade arrays and non-rotating arrays. Therefore, some form of sealing mechanism is required to discourage fluid from flowing radially inward between the adjacent arrays. In addition to the loss of efficiency because of fluid escaping around the arrays of blades, gas turbine engine components located radially inward of the flow path may be damaged by contact with the hot gases from the flow path. Such components include rotor disks, which are under significant stress. As is well known, increasing the operating temperature of the rotor disk decreases the allowable stress of the disk material.
One popular form of sealing mechanism is a knife edge element engaged with a honeycomb type structure. Typically, the knife edge is extended from the rotating component and the honeycomb material is attached to the non-rotating component. The honeycomb material is formed from very thin (on the order of 0.004 in) sheet metal in the shape of open cells. During operation, the knife edge may engage the honeycomb material and wear a groove into the honeycomb material. The wearing of the honeycomb accounts for tolerances between the components and for thermal growth during operation. This type of sealing arrangement is desirable because the honeycomb material is inexpensive and is generally easily replaced once it wears away.
A drawback to using honeycomb material in a sealing mechanism is that it quickly degrades if exposed to the high temperatures present in the fluid flowing through the flow path. Degradation due to heat exposure causes the honeycomb seal to be replaced prematurely, i.e. prior to wearing out due to engagement with the knife edge. To account for this, honeycomb seals used in hot sections of the gas turbine engine are coated with a thermal barrier coating (TBC). The TBC protects the outward facing surfaces of the honeycomb. Unfortunately, the TBC applied to the honeycomb is often different from the TBC applied to the airfoil because the sheet metal of the honeycomb cannot withstand the high temperatures associated with the processes required to apply the common TBC used on airfoils. The added expense of a unique TBC and the expense of an additional step to apply the TBC increases the cost of fabricating the airfoil. Further, since the honeycomb seals are frequently replaced during the life of the airfoil, the costs associated with repairing and maintaining the airfoil may be excessive.
The above art notwithstanding, scientists and engineers under the direction of Applicants' Assignee are working to develop turbine components, such as airfoils, that have longer operational life expectancies and that are inexpensive to maintain.